Aircraft control mode transition smoothing

ABSTRACT

In accordance with an embodiment, a method of operating an aircraft includes operating the aircraft in a first mode including determining an attitude based on a pilot stick signal, where a translational speed or an attitude of the aircraft is proportional to an amplitude of the pilot stick signal in the first mode; transitioning from the first mode to a second mode when a velocity of the aircraft exceeds a first velocity threshold; and operating the aircraft in the second mode where the output of the rate controller is proportional to the amplitude of the pilot stick signal.

This application is a continuation of U.S. patent application Ser. No.17/346,989, filed Jun. 14, 2021, which application is a continuation ofU.S. patent application Ser. No. 16/253,037, filed Jan. 21, 2019, nowU.S. Pat. No. 11,067,981, which application is a continuation in part ofU.S. patent application Ser. No. 16/108,479, filed on Aug. 22, 2018, nowU.S. Pat. No. 10,691,140, and U.S. patent application Ser. No.15/446,911, filed on Mar. 1, 2017, now U.S. Pat. No. 10,481,615, whichapplications are hereby incorporated herein by reference in theirentireties.

TECHNICAL FIELD

The present invention relates generally to a system and method for aflight control, and, in particular embodiments, to a system and methodfor control mode transition smoothing for an aircraft.

BACKGROUND

Fly-by-wire systems in aircraft, as opposed to mechanically controlledsystems, use electronic signals to control the flight surfaces andengines in the aircraft. For example, instead of having the pilotcontrols mechanically linked to the control surfaces via a hydraulicsystem, the pilot controls are electronically linked to a flightcomputer, which, in turn, controls flight surface actuators viaelectronic signals. By further interfacing the flight computer toaircraft sensors, sophisticated control algorithms may be used toprovide autopilot functionality, as well as to stabilize and control theaircraft.

While fly-by-wire systems have become commonplace in commercial andcivilian fixed wing aircraft, their adoption among rotorcraft, such ashelicopters, has been much slower due, in part, to the increasedcomplexity of controlling and stabilizing a rotorcraft. However, byadopting fly-by-wire systems in helicopters, safer operation may beachieved in difficult flight environments such as low speed, lowaltitude, degraded visual environments and inclement weather. Anotherarea in which fly-by-wire systems may benefit rotorcraft is in thereduction in pilot workload. By providing automated features such asstabilization in response to wind, control axis decoupling, positionhold and heading hold functionality, the pilot is freed up to focus onthe environment in which he flies.

One challenge in the design of fly-by-wire systems for rotorcraft istransitioning between different modes of operation that utilizedifferent control laws or algorithms. In some circumstances, the changein control algorithm may result in a physical transient during operationof the rotorcraft that might be physically discernable as a bump or joltby the pilot or passengers.

SUMMARY

In accordance with an embodiment, a method of operating an aircraftincludes operating the aircraft in a first mode including determining anattitude based on a pilot stick signal generated by a pilot stickassembly, determining a first rate command based on the determinedattitude using an attitude controller, determining an actuator commandbased on the determined first rate command, determining the actuatorcommand including using a rate controller having an integrator, andproviding an output of the rate controller to an actuator, where atranslational speed or an attitude of the aircraft is proportional to anamplitude of the pilot stick signal in the first mode; transitioningfrom the first mode to a second mode when a velocity of the aircraftexceeds a first velocity threshold, transitioning including fading out again of the attitude controller over a first period of time; andoperating the aircraft in the second mode including providing the pilotstick signal to an input of the rate controller, where the output of therate controller is proportional to the amplitude of the pilot sticksignal.

In accordance with another embodiment, a flight control system for anaircraft includes: a processor and a non-transitory computer readablestorage medium with an executable program stored thereon, the executableprogram including instructions to: receive a pilot control signal via afirst interface of the processor; in a first mode determine an attitudebased on the received pilot control signal, determine a first ratecommand based on the determined attitude using an attitude controller,determine an actuator command based on the determined first ratecommand, determining the actuator command including executing a ratecontroller that has an integrator, and providing an output of the ratecontroller to an actuator via a second interface of the processor, wherea state of the aircraft corresponding to the attitude is configured tobe proportional to the received pilot control signal; and transitioningfrom the first mode to a second mode when a first condition of theaircraft crosses a first predetermined threshold, transitioningincluding fading out a gain of the attitude controller over a firstperiod of time, and decreasing a value of the integrator over a secondperiod of time; and in the second mode, providing the pilot controlsignal to an input of the rate controller, where the output of the ratecontroller is proportional to the received pilot control signal.

In accordance with a further embodiment, a rotorcraft includes: a body;a power train coupled to the body and including a power source and adrive shaft coupled to the power source; a rotor system coupled to thepower train and including a plurality of rotor blades; a flight controlsystem operable to change at least one operating condition of the rotorsystem; a pilot control assembly configured to receive commands from apilot, where the flight control system is a fly-by-wire flight controlsystem in electrical communication with the pilot control assembly; anda flight control computer in electrical communication between the flightcontrol system and the pilot control assembly, the flight controlcomputer configured to: receive, from the pilot control assembly a pilotcommand to change a first flight characteristic, when a velocity of therotorcraft is less than a first velocity threshold, interpret the firstflight characteristic as a requested translational speed or as arequested attitude in a first mode, determine a controlled attitudebased on the requested translational speed or the requested attitudeusing an attitude controller, determining a first rate command based onthe determined attitude, and determine an actuator command based on thedetermined rate command using a rate controller; when the velocity ofthe rotorcraft is greater than a second velocity threshold, interpretthe first flight characteristic as a rate in a second mode by providingthe pilot command to the rate controller, and when the velocity of therotorcraft increases past the second velocity threshold, fade out a gainof the attitude controller, and successively decrease a value of anintegrator in the rate controller.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the present invention, and theadvantages thereof, reference is now made to the following descriptionstaken in conjunction with the accompanying drawings, in which:

FIG. 1 illustrates an embodiment rotorcraft;

FIG. 2 illustrates a block diagram of an embodiment rotorcraft flightcontrol system;

FIG. 3 illustrates a block diagram of an embodiment flight controlsystem;

FIG. 4 illustrates a block diagram of a further embodiment flightcontrol system;

FIG. 5 illustrates a block diagram of an embodiment TRC mode attitudecontroller;

FIG. 6 illustrates a block diagram of an embodiment method;

FIG. 7 illustrates an embodiment computer system;

FIG. 8A is a table illustrating embodiment flight modes; and FIG. 8B isa table illustrating embodiment mode transitions;

FIG. 9 illustrates a block diagram of an embodiment flight controlsystem that supports multiple mode transitions;

FIG. 10 illustrates a block diagram of an embodiment attitudecontroller;

FIG. 11 illustrates a block diagram of an embodiment method of changingflight control modes;

FIG. 12 illustrates a block diagram of a flight control system thatsupports multiple mode transitions according to a further embodiment;

FIG. 13 illustrates a block diagram of an attitude controller accordingto a further embodiment;

FIGS. 14A-14B illustrate block diagrams of embodiment rate controllers;and

FIG. 15 illustrates a block diagram a method of changing flight controlmodes according to a further embodiment.

Corresponding numerals and symbols in different figures generally referto corresponding parts unless otherwise indicated. The figures are drawnto clearly illustrate the relevant aspects of the preferred embodimentsand are not necessarily drawn to scale. To more clearly illustratecertain embodiments, a letter indicating variations of the samestructure, material, or process step may follow a figure number.

DETAILED DESCRIPTION OF ILLUSTRATIVE EMBODIMENTS

Illustrative embodiments of the system and method of the presentdisclosure are described below. In the interest of clarity, all featuresof an actual implementation may not be described in this specification.It will of course be appreciated that in the development of any suchactual embodiment, numerous implementation-specific decisions may bemade to achieve the developer's specific goals, such as compliance withsystem-related and business-related constraints, which will vary fromone implementation to another. Moreover, it should be appreciated thatsuch a development effort might be complex and time-consuming but wouldnevertheless be a routine undertaking for those of ordinary skill in theart having the benefit of this disclosure.

Reference may be made herein to the spatial relationships betweenvarious components and to the spatial orientation of various aspects ofcomponents as the devices are depicted in the attached drawings.However, as will be recognized by those skilled in the art after acomplete reading of the present disclosure, the devices, members,apparatuses, etc. described herein may be positioned in any desiredorientation. Thus, the use of terms such as “above,” “below,” “upper,”“lower,” or other like terms to describe a spatial relationship betweenvarious components or to describe the spatial orientation of aspects ofsuch components should be understood to describe a relative relationshipbetween the components or a spatial orientation of aspects of suchcomponents, respectively, as the device described herein may be orientedin any desired direction.

The increasing use of rotorcraft, in particular, for commercial andindustrial applications, has led to the development of larger morecomplex rotorcraft. However, as rotorcraft become larger and morecomplex, the differences between flying rotorcraft and fixed wingaircraft has become more pronounced. Since rotorcraft use one or moremain rotors to simultaneously provide lift, control attitude, controlaltitude, and provide lateral or positional movement, different flightparameters and controls are tightly coupled to each other, as theaerodynamic characteristics of the main rotors affect each control andmovement axis. For example, the flight characteristics of a rotorcraftat cruising speed or high speed may be significantly different than theflight characteristics at hover or at relatively low speeds.Additionally, different flight control inputs for different axes on themain rotor, such as cyclic inputs or collective inputs, affect otherflight controls or flight characteristics of the rotorcraft. Forexample, pitching the nose of a rotorcraft forward to increase forwardspeed will generally cause the rotorcraft to lose altitude. In such asituation, the collective may be increased to maintain level flight, butthe increase in collective causes increased power to the main rotorwhich, in turn, requires additional anti-torque force from the tailrotor. This is in contrast to fixed wing systems where the controlinputs are less closely tied to each other and flight characteristics indifferent speed regimes are more closely related to each other.

Recently, fly-by-wire (FBW) systems have been introduced in rotorcraftto assist pilots in stably flying the rotorcraft and to reduce workloadon the pilots. The FBW system may provide different controlcharacteristics or responses for cyclic, pedal or collective controlinput in the different flight regimes, and may provide stabilityassistance or enhancement by decoupling physical flight characteristicsso that a pilot is relieved from needing to compensate for some flightcommands issued to the rotorcraft. FBW systems may be implemented in oneor more flight control computers (FCCs) disposed between the pilotcontrols and flight control systems, providing corrections to flightcontrols that assist in operating the rotorcraft more efficiently orthat put the rotorcraft into a stable flight mode while still allowingthe pilot to override the FBW control inputs. The FBW systems in arotorcraft may, for example, automatically adjust power output by theengine to match a collective control input, apply collective or powercorrection during a cyclic control input, provide automation of one ormore flight control procedures, provide for default or suggested controlpositioning, or the like.

FBW systems for rotorcraft must provide stable flight characteristicsfor FBW controlled flight parameters while permitting the pilot tooverride or adjust any suggested flight parameters suggested by the FBWsystem. Additionally, in providing enhanced control and automatedfunctionality for rotorcraft flight, the FBW system must maintain anintuitive and easy to use flight control system for the pilot. Thus, theFBW system adjusts the pilot flight controls so that the controls are ina position associated with the relevant flight parameter. For example,the FBW system may adjust the collective stick to provide suggested orFBW controlled flight parameters, and which reflect a collective orpower setting. Thus, when the pilot releases the collective stick andthe FBW system provides collective control commands, the collectivestick is positioned intuitively in relation to the actual power orcollective setting so that, when the pilot grasps the collective stickto retake control, the control stick is positioned where the pilotexpects the stick to be positioned for the actual collective setting ofthe main rotor. Similarly, the FBW system use the cyclic stick to, forexample, adjust for turbulence, drift or other disturbance to the flightpath, and may move the cyclic stick as the FBW system compensates thecyclic control. Thus, when the pilot grasps the cyclic stick to takecontrol of flight from the FBW system, the cyclic stick is positioned toreflect the actual cyclic settings.

Embodiments of the present disclosure will be described with respect topreferred embodiments in a specific context, namely a system and methodfor smoothing a transition between a translational rate command (TRC)mode to a rate mode in a rotorcraft. Embodiments of the presentdisclosure may also be applied to other control mode transitions in theoperation and control of a rotorcraft.

In an embodiment of the present invention, a forward displacement on thecyclic controller or pilot stick of the rotorcraft is interpreted in oneof two ways. When the speed of the rotorcraft is 10 kts or less and thespeed command by displacement on the cyclic controller is less than 10kts, the displacement on the cyclic controller is interpreted as a speedcommand. For example, ½ inch of stick displacement may correspond to aspeed of 2 kts; 1″ of stick displacement may correspond to a speed of 4kts, and so on. However, when the speed of the rotorcraft is greaterthan 10 kts or the speed command by displacement on the cycliccontroller is greater than 10 kts, displacement on the cyclic controlleris interpreted as a rate command. For example, ½″ of stick displacementmay correspond to a rate of 10° per second; 1″ of stick displacement maycorrespond to a rate of 20° per second, and so on. By automaticallyswitching between the two different modes of operation based on thespeed of the rotorcraft, pilot workload may be reduced.

In some circumstances, however, the rotorcraft may undergo a physicaltransient during the transition between the speed control mode at lowspeeds and the rate control mode at higher speeds. This physicaltransient may be caused by sudden changes in flight control algorithmsand abrupt shifts in control signals. In embodiments of the presentinvention, the transition between speed control and rate control is madesmoother by controlling the gains and the signal path structure of anattitude controller used during the speed control mode. In one example,the overall gain of the attitude controller is reduced using a fader.During this time, a proportional path in the controller is maintainedwhile a value of an integrator is decremented to zero. Also during thistime the gain of a control path between the cyclic controller and therate control block is increased using a fader. By maintaining aproportional path in the attitude controller, decrementing theintegrator to zero, fading out the overall gain of the attitudecontroller and fading in the gain of the control path between the cycliccontroller and the rate control block, a smooth transition between speedcontrol mode and rate control mode may be made more smooth. It should beunderstood that embodiments of the present invention may be applied toother control signals and control paths of the rotorcraft.

FIG. 1 illustrates a rotorcraft 101 according to some embodiments. Therotorcraft 101 has a main rotor system 103, which includes a pluralityof main rotor blades 105. The pitch of each main rotor blade 105 may becontrolled by a swashplate 107 in order to selectively control theattitude, altitude and movement of the rotorcraft 101. The swashplate107 may be used to collectively and/or cyclically change the pitch ofthe main rotor blades 105. The rotorcraft 101 also has an anti-torquesystem, which may include a tail rotor 109, no-tail-rotor (NOTAR), ordual main rotor system. In rotorcraft with a tail rotor 109, the pitchof each tail rotor blade 111 is collectively changed in order to varythrust of the anti-torque system, providing directional control of therotorcraft 101. The pitch of the tail rotor blades 111 is changed by oneor more tail rotor actuators. In some embodiments, the FBW system sendselectrical signals to the tail rotor actuators or main rotor actuatorsto control flight of the rotorcraft.

Power is supplied to the main rotor system 103 and the anti-torquesystem by engines 115. There may be one or more engines 115, which maybe controlled according to signals from the FBW system. The output ofthe engine 115 is provided to a driveshaft 117, which is mechanicallyand operatively coupled to the rotor system 103 and the anti-torquesystem through a main rotor transmission 119 and a tail rotortransmission, respectively.

The rotorcraft 101 further includes a fuselage 125 and tail section 123.The tail section 123 may have other flight control devices such ashorizontal or vertical stabilizers, rudder, elevators, or other controlor stabilizing surfaces that are used to control or stabilize flight ofthe rotorcraft 101. The fuselage 125 includes a cockpit 127, whichincludes displays, controls, and instruments. It should be appreciatedthat even though rotorcraft 101 is depicted as having certainillustrated features, the rotorcraft 101 may have a variety ofimplementation-specific configurations. For instance, in someembodiments, cockpit 127 is configured to accommodate a pilot or a pilotand co-pilot, as illustrated. It is also contemplated, however, thatrotorcraft 101 may be operated remotely, in which case cockpit 127 couldbe configured as a fully functioning cockpit to accommodate a pilot (andpossibly a co-pilot as well) to provide for greater flexibility of use,or could be configured with a cockpit having limited functionality(e.g., a cockpit with accommodations for only one person who wouldfunction as the pilot operating perhaps with a remote co-pilot or whowould function as a co-pilot or back-up pilot with the primary pilotingfunctions being performed remotely). In yet other contemplatedembodiments, rotorcraft 101 could be configured as an unmanned vehicle,in which case cockpit 127 could be eliminated entirely in order to savespace and cost.

FIG. 2 illustrates a fly-by-wire flight control system 201 for arotorcraft according to some embodiments. A pilot may manipulate one ormore pilot flight controls in order to control flight of the rotorcraft.The pilot flight controls may include manual controls such as a cyclicstick 231 in a cyclic control assembly 217, a collective stick 233 in acollective control assembly 219, and pedals 239 in a pedal assembly 221.Inputs provided by the pilot to the pilot flight controls may betransmitted mechanically and/or electronically (e.g., via the FBW flightcontrol system) to flight control devices by the flight control system201. Flight control devices may represent devices operable to change theflight characteristics of the rotorcraft. Flight control devices on therotorcraft may include mechanical and/or electrical systems operable tochange the positions or angle of attack of the main rotor blades 105 andthe tail rotor blades in or to change the power output of the engines115, as examples. Flight control devices include systems such as theswashplate 107, tail rotor actuator 113, and systems operable to controlthe engines 115. The flight control system 201 may adjust the flightcontrol devices independently of the flight crew in order to stabilizethe rotorcraft, reduce workload of the flight crew, and the like. Theflight control system 201 includes engine control computers (ECCUs) 203,flight control computers 205, and aircraft sensors 207, whichcollectively adjust the flight control devices.

The flight control system 201 has one or more flight control computers205 (FCCs). In some embodiments, multiple FCCs 205 are provided forredundancy. One or more modules within the FCCs 205 may be partially orwholly embodied as software and/or hardware for performing anyfunctionality described herein. In embodiments where the flight controlsystem 201 is a FBW flight control system, the FCCs 205 may analyzepilot inputs and dispatch corresponding commands to the ECCUs 203, thetail rotor actuator 113, and/or actuators for the swashplate 107.Further, the FCCs 205 are configured and receive input commands from thepilot controls through sensors associated with each of the pilot flightcontrols. The input commands are received by measuring the positions ofthe pilot controls. The FCCs 205 also control tactile cueing commands tothe pilot controls or display information in instruments on, forexample, an instrument panel 241.

The ECCUs 203 control the engines 115. For example, the ECCUs 203 mayvary the output power of the engines 115 to control the rotational speedof the main rotor blades or the tail rotor blades. The ECCUs 203 maycontrol the output power of the engines 115 according to commands fromthe FCCs 205, or may do so based on feedback such a measured revolutionsper minute (RPM) of the main rotor blades.

The aircraft sensors 207 are in communication with the FCCs 205. Theaircraft sensors 207 may include sensors for measuring a variety ofrotorcraft systems, flight parameters, environmental conditions and thelike. For example, the aircraft sensors 207 may include sensors formeasuring airspeed, altitude, attitude, position, orientation,temperature, airspeed, vertical speed, and the like. Other sensors 207could include sensors relying upon data or signals originating externalto the rotorcraft, such as a global positioning system (GPS) sensor, aVHF Omnidirectional Range sensor, Instrument Landing System (ILS), andthe like.

The cyclic control assembly 217 is connected to a cyclic trim assembly229 having one or more cyclic position sensors 211, one or more cyclicdetent sensors 235, and one or more cyclic actuators or cyclic trimmotors 209. The cyclic position sensors 211 measure the position of thecyclic control stick 231. In some embodiments, the cyclic control stick231 is a single control stick that moves along two axes and permits apilot to control pitch, which is the vertical angle of the nose of therotorcraft and roll, which is the side-to-side angle of the rotorcraft.In some embodiments, the cyclic control assembly 217 has separate cyclicposition sensors 211 that measuring roll and pitch separately. Thecyclic position sensors 211 for detecting roll and pitch generate rolland pitch signals, respectively, (sometimes referred to as cycliclongitude and cyclic latitude signals, respectively) which are sent tothe FCCs 205, which controls the swashplate 107, engines 115, tail rotor109 or related flight control devices.

The cyclic trim motors 209 are connected to the FCCs 205, and receivesignals from the FCCs 205 to move the cyclic control stick 231. In someembodiments, the FCCs 205 determine a suggested cyclic stick positionfor the cyclic stick 231 according to one or more of the collectivestick position, the pedal position, the speed, altitude and attitude ofthe rotorcraft, the engine revolutions per minute (RPM), enginetemperature, main rotor RPM, engine torque or other rotorcraft systemconditions or flight conditions. The suggested cyclic stick position isa positon determined by the FCCs 205 to give a desired cyclic action. Insome embodiments, the FCCs 205 send a suggested cyclic stick positionsignal indicating the suggested cyclic stick position to the cyclic trimmotors 209. While the FCCs 205 may command the cyclic trim motors 209 tomove the cyclic stick 231 to a particular position (which would in turndrive actuators associated with swashplate 107 accordingly), the cyclicposition sensors 211 detect the actual position of the cyclic stick 231that is set by the cyclic trim motors 206 or input by the pilot,allowing the pilot to override the suggested cyclic stick position. Thecyclic trim motor 209 is connected to the cyclic stick 231 so that thepilot may move the cyclic stick 231 while the trim motor is driving thecyclic stick 231 to override the suggested cyclic stick position. Thus,in some embodiments, the FCCs 205 receive a signal from the cyclicposition sensors 211 indicating the actual cyclic stick position, and donot rely on the suggested cyclic stick position to command theswashplate 107.

Similar to the cyclic control assembly 217, the collective controlassembly 219 is connected to a collective trim assembly 225 having oneor more collective position sensors 215, one or more collective detentsensors 237, and one or more collective actuators or collective trimmotors 213. The collective position sensors 215 measure the position ofa collective control stick 233 in the collective control assembly 219.In some embodiments, the collective control stick 233 is a singlecontrol stick that moves along a single axis or with a lever typeaction. A collective position sensor 215 detects the position of thecollective control stick 233 and sends a collective position signal tothe FCCs 205, which controls engines 115, swashplate actuators, orrelated flight control devices according to the collective positionsignal to control the vertical movement of the rotorcraft. In someembodiments, the FCCs 205 may send a power command signal to the ECCUs203 and a collective command signal to the main rotor or swashplateactuators so that the angle of attack of the main blades is raised orlowered collectively, and the engine power is set to provide the neededpower to keep the main rotor RPM substantially constant.

The collective trim motor 213 is connected to the FCCs 205, and receivessignals from the FCCs 205 to move the collective control stick 233.Similar to the determination of the suggested cyclic stick position, insome embodiments, the FCCs 205 determine a suggested collective stickposition for the collective control stick 233 according to one or moreof the cyclic stick position, the pedal position, the speed, altitudeand attitude of the rotorcraft, the engine RPM, engine temperature, mainrotor RPM, engine torque or other rotorcraft system conditions or flightconditions. The FCCs 205 generate the suggested collective stickposition and send a corresponding suggested collective stick signal tothe collective trim motors 213 to move the collective stick 233 to aparticular position. The collective position sensors 215 detect theactual position of the collective stick 233 that is set by thecollective trim motor 213 or input by the pilot, allowing the pilot tooverride the suggested collective stick position.

The pedal control assembly 221 has one or more pedal sensors 227 thatmeasure the position of pedals or other input elements in the pedalcontrol assembly 221. In some embodiments, the pedal control assembly221 is free of a trim motor or actuator, and may have a mechanicalreturn element that centers the pedals when the pilot releases thepedals. In other embodiments, the pedal control assembly 221 has one ormore trim motors that drive the pedal to a suggested pedal positionaccording to a signal from the FCCs 205. The pedal sensor 227 detectsthe position of the pedals 239 and sends a pedal position signal to theFCCs 205, which controls the tail rotor 109 to cause the rotorcraft toyaw or rotate around a vertical axis.

The cyclic and collective trim motors 209 and 213 may drive the cyclicstick 231 and collective stick 233, respectively, to suggestedpositions. The cyclic and collective trim motors 209 and 213 may drivethe cyclic stick 231 and collective stick 233, respectively, tosuggested positions, but this movement capability may also be used toprovide tactile cueing to a pilot. The trim motors 209 and 213 may pushthe respective stick in a particular direction when the pilot is movingthe stick to indicate a particular condition. Since the FBW systemmechanically disconnects the stick from one or more flight controldevices, a pilot may not feel a hard stop, vibration, or other tactilecue that would be inherent in a stick that is mechanically connected toa flight control assembly. In some embodiments, the FCCs 205 may causethe trim motors 209 and 213 to push against a pilot command so that thepilot feels a resistive force, or may command one or more frictiondevices to provide friction that is felt when the pilot moves the stick.Thus, the FCCs 205 control the feel of a stick by providing pressureand/or friction on the stick.

Additionally, the cyclic control assembly 217, collective controlassembly 219 and/or pedal control assembly 221 may each have one or moredetent sensors that determine whether the pilot is handling a particularcontrol device. For example, the cyclic control assembly 217 may have acyclic detent sensor 235 that determines that the pilot is eitherresting or moving the cyclic stick 231, while the collective controlassembly 219 has a collective detent sensor 237 that determines whetherthe pilot is either holding or moving the collective stick 233. Thesedetent sensors 235, 237 detect motion and/or position of the respectivecontrol stick that is caused by pilot input, as opposed to motion and/orposition caused by commands from the FCCs 205, rotorcraft vibration, andthe like and provide feedback signals indicative of such to the FCCs.When the FCCs 205 detect that a pilot has control of, or ismanipulating, a particular control, the FCCs 205 may determine thatstick to be out-of-detent (00D). Likewise, the FCCs may determine thatthe stick is in-detent (ID) when the signals from the detent sensorsindicate to the FCCs 205 that the pilot has released a particular stick.The FCCs 205 may provide different default control or automated commandsto one or more flight systems based on the detent status of a particularstick or pilot control.

Moving now to the operational aspects of flight control system 201, FIG.3 illustrates in a highly schematic fashion, a manner in which flightcontrol system 201 may implement FBW functions as a series ofinter-related feedback loops running certain control laws. FIG. 3representatively illustrates a three-loop flight control system 201according to an embodiment. In some embodiments, elements of thethree-loop flight control system 201 may be implemented at leastpartially by FCCs 205. As shown in FIG. 3 , however, all, some, or noneof the components (301, 303, 305, 307) of three-loop flight controlsystem 201 could be located external or remote from the rotorcraft 100and communicate to on-board devices through a network connection 309.

The three-loop flight control system 201 of FIG. 3 has a pilot input311, an outer loop 313, a rate (middle) loop 315, an inner loop 317, adecoupler 319, and aircraft equipment 321 (corresponding, e.g., toflight control devices such as swashplate 107, tail rotor transmission212, etc., to actuators (not shown) driving the flight control devices,to sensors such as aircraft sensors 207, position sensors 211, 215,detent sensors 235, 237, etc., and the like).

In the example of FIG. 3 , a three-loop design separates the innerstabilization and rate feedback loops from outer guidance and trackingloops. The control law structure primarily assigns the overallstabilization task and related tasks of reducing pilot workload to innerloop 317. Next, middle loop 315 provides rate augmentation. Outer loop313 focuses on guidance and tracking tasks. Since inner loop 317 andrate loop 315 provide most of the stabilization, less control effort isrequired at the outer loop level. As representatively illustrated inFIG. 3 , a switch 322 may be provided to turn outer loop flightaugmentation on and off, as the tasks of outer loop 313 are notnecessary for flight stabilization.

In some embodiments, the inner loop 317 and rate loop 315 include a setof gains and filters applied to roll/pitch/yaw 3-axis rate gyro andacceleration feedback sensors. Both the inner loop and rate loop maystay active, independent of various outer loop hold modes. Outer loop313 may include cascaded layers of loops, including an attitude loop, aspeed loop, a position loop, a vertical speed loop, an altitude loop,and a heading loop. In accordance with some embodiments, the controllaws running in the illustrated the loops allow for decoupling ofotherwise coupled flight characteristics, which in turn may provide formore stable flight characteristics and reduced pilot workload.Furthermore, the outer loop 313 may allow for automated orsemi-automated operation of certain high-level tasks or flight patterns,thus further relieving the pilot workload and allowing the pilot tofocus on other matters including observation of the surrounding terrain.

FIG. 4 illustrates a flight control system 400 according to anembodiment of the present invention. Pilot stick pitch block 401 a andpilot stick roll block 401 b represent, for example, the respectivepitch and roll commands emanating from cyclic controller 217 of therotorcraft. As shown, pilot stick blocks 401 a and 401 b interface toflight controller 402. In various embodiments, flight controller 402 isimplemented using a flight computer, or other processing hardware.Flight controller 402 also interfaces with and controls aircraftequipment 321 representing various actuators, sensors, and the physicalbody of the rotorcraft. In various embodiments, flight controller 402controls aircraft equipment 321 using three loops: an inner loop, a ratefeedback loop and a state feedback loop: the inner loop stabilizes thedynamics of the rotorcraft, the rate loop controls the angular rates ofthe rotorcraft, and the outer loop provides control signals to the innerloop and/or rate loops to effect a desired attitude, speed and positionof the rotorcraft. In some embodiments, the outer loop supports andprovides flight augmentation or auto-pilot functionality and may bemanually or automatically disabled based on flight and systemconditions. The inner loop and rate feedback loops, on the other hand,remain operational to provide stability to the rotorcraft.

For purposes of illustration, flight controller 402 is illustrated withrespect to the general control blocks that affect the rotational rate ofan embodiment rotorcraft, namely the control blocks affecting the rollrate and the pitch rate of the rotorcraft. It should be understood thatflight controller 402 may also include other controllers and controlpaths that affect the yaw and other states of the rotorcraft in additionto the pitch rate and roll rate. As shown, the inner stabilization loopis controlled by inner loop controller 317, the rate loop is controlledby pitch rate controller 315 a and roll rate controller 315 b, and theouter loop is controlled by outer loop controller 313 in conjunctionwith TRC mode attitude controllers 408 a and 408 b that provide TRC modepitch and roll control, respectively, during embodiment TRC modes.

Each of inner loop controller 317, decoupler 319 and rate controller 315may be implemented using flight control algorithms known in the art.Inner loop controller 317 receives sensor feedback from sensors such asgyroscopes and accelerometers within the rotorcraft and provides controlsignals to various actuators, such as swashplate 107 to stabilize therotorcraft. Rate controller 315 receives rate feedback from rategyroscopes on all axes and provides a rate command signal to inner loopcontroller 317 based on the rate feedback and the position of pilotstick pitch block 401 a and pilot stick roll block 401 b in some modesof operation. Decoupler 319 receives the various rate commands andapproximately decouples all 4-axes (pitch, roll, yaw, and vertical) suchthat, for example, a forward longitudinal stick input does not requirethe pilot to push the stick diagonally.

V_(T)=k√{square root over (S_(LON) ²+S_(LAT) ²)} Outer loop controller313 receives state feedback from the sensors of aircraft equipment 321.This state feedback may include, for example, speed, position andattitude. In a translational rate commend (TRC) mode, outer loopcontroller 313 receives a command from the pilot stick represented bypilot stick pitch block 401 a and pilot stick roll block 401 b,determines a corresponding translational speed based on the pilot stickcommand, and determines corresponding pitch and roll attitude commandsbased on the determined translational speed. The translational speed mayinclude a forward component based on commands from pilot stick pitchblock 401 a and a lateral component based on commands from pilot stickroll block 401 b. In some embodiments, the corresponding translationalspeed is determined by multiplying the pilot stick command, or a vectorsum of the pitch and roll pilot stick commands with a scale factorand/or by using a lookup table. The translational speed may becalculated using a vector sum as follows:V _(T) =k√{square root over (S _(LON) ² +S _(LAT) ²)},

V_(T)=k√{square root over (S_(LON) ²+S_(LAT) ² where V_(T))} is thetranslational speed, k is a scale factor, S_(LON) is the pilot stickpitch command, and S_(LAT) is the pilot stick roll command. In someembodiments, the vector sum is not used and the translational speed isbased on the forward speed and/or the lateral speed.

The pitch and roll attitude commands may be determined from thetranslational speed using a speed control loop. During the TRC mode, TRCmode attitude controllers 408 a and 408 b individually calculate pitchand roll attitude errors by subtracting the pitch attitude feedback fromthe pitch attitude command and by subtracting the roll attitude feedbackfrom the roll attitude command. In various embodiments, the pitch androll attitude feedback is a component of the state feedback. TRC modeattitude controller 408 a applies a dynamic control algorithm to thepitch attitude error to produce an outer loop pitch rate command, andTRC mode attitude controller 408 b applies a dynamic control algorithmto the roll attitude error to produce an outer loop roll rate command.These outer loop rate commands are applied to decoupler 319 via summingblocks 412 a and 412 b. While the embodiment of FIG. 4 shows TRC modeattitude controllers 408 a and 408 b used to provide TRC mode control toboth the pitch and roll channels, it should be understood that inalternative embodiments, TRC mode control may be applied to a singleattitude channel, such as the pitch channel for speed control in theforward direction.

In various embodiments of the present invention, the TRC mode isautomatically selected based on the velocity or ground speed of therotorcraft. This ground speed may be measured, for example, using GPS oran optical sensing system such as laser-based sensing system. In oneexample, the TRC mode is selected when the ground speed is less than 10kts, in which a fixed offset of the pilot stick effects a constanttranslation. When the ground speed is 10 kts and greater, the pilotstick effects a rate in a rate control mode. In the embodiment of FIG. 4, the TRC mode is activated by enabling TRC mode attitude controller 408a and 408 b when the ground speed is less than 10 kts and decoupling thepilot stick pitch block 401 a from pitch rate controller 315 a via fader404 a and optionally decoupling the pilot stick roll block 401 b fromthe roll rate controller 315 b via fader 404 b. It should be appreciatedthat the 10 kts threshold is just one example of many possiblegroundspeed thresholds. In alternative embodiments, other thresholdvalues may be used. Likewise, the rate control mode is selected bycoupling pilot stick pitch block 401 a to pitch rate controller 315 avia fader 404 a, and by coupling pilot stick roll block 401 b to rollrate controller 315 b via fader 404 b. In some embodiments, pilot stickblocks 401 a and 401 b are coupled to rate controllers 315 a and 315 bwhen the pilot stick is out of detent. In some embodiments, hysteresismay be applied to the ground speed threshold to prevent metastability ofthe control mode when the rotorcraft is flying at about 10 kts. In suchan embodiment the ground speed threshold for transitioning from the TRCto the rate control mode is greater than the ground speed threshold fortransitioning from the rate control mode to the TRC mode. For example,in one embodiment, TRC mode attitude controllers 408 a and 408 b aredisabled when the pilot stick command is greater than 10 kts and thenre-enabled when the pilot stick command is less than 3 kts.Alternatively, other hysteresis/threshold values may be used.Alternatively, a single threshold may be used without hysteresis. Invarious embodiments, the 10 kts and 3 kts speed thresholds may beapplied to the forward speed, the lateral speed or a combination of theforward and lateral speed. Such a combination may be determined, forexample, by calculating a vector sum of the forward speed and thelateral speed.

During operation, faders 404 a and 404 b provide a gain of one when thepilot stick is out of detent, and provides a gain of zero when the pilotstick is in detent. In various embodiments, the gain of faders 404 a and404 b linearly increase or decrease between zero and one over apredetermined period of time or change in a predetermined speed range.This predetermined period of time is between about 1 s and about 10 s;however, fading times outside of this range may be implemented dependingon the particular embodiment and its specifications. In furtheralternative embodiments, the gains of fader 404 a and 404 b may bedifferent from one and zero. In some embodiments, pitch rate controllers315 a and roll rate controller 315 b each a high-pass filter and/orwashout filter coupled to their respective inputs. Thus, when the stickis out of detent and held at a steady displacement, the effectivecommand produced by rate controller 315 a and 315 b as a result of apilot stick command goes to zero. This high-pass filter or washoutfilter also prevents the output of TRC mode attitude controller 408 aand 408 b and the output of faders 404 a and 404 b from conflicting or“fighting” with each other when the pilot stick is out of detent.

FIG. 5 illustrates a block diagram of TRC mode attitude controller 408according to an embodiment of the present invention that may be used toimplement TRC attitude controllers 408 a and 408 b illustrated in FIG. 4. As shown, TRC mode attitude controller 408 includes subtractor 430,fader 432, dynamic compensation block 434 and proportional-integral (PI)controller 450. Subtractor 430 produces an error signal based on theattitude command produced by the outer loop controller 313 (FIG. 4 ) anddynamic compensation block 434 compensates for the dynamics of therotorcraft in order to improve stability and/or adjust the time responseof the attitude loop. Dynamic compensation block 434 may include variouscontrol blocks known in the art including but not limited a PIDcontroller and a lead-lag compensator.

PI controller 450 includes a proportional signal path that includesproportional gain block 436 coupled in parallel with an integral signalpath that includes integral gain block 438 and integrator 440. Invarious embodiments, proportional gain block 436 and integral gain block438 may be implemented, for example, by performing a multiplication orscaling operation. Integrator 440 may be implemented, for example, usingan accumulator.

During TRC mode operation below the 10 kts speed threshold, fader 432has a gain of one and the input of integrator 440 is coupled to theoutput of dynamic compensation block 434 via integral gain block 438. Onthe other hand, in the rate control mode in which the rotorcraft has aground speed of 10 kts and above or the speed command by displacement onthe cyclic controller is greater than 10 kts, fader 432 has a gain ofzero, and the input of integrator 440 is coupled to its output viafeedback gain block 444 and limiter 442, which effectively causesintegrator 440 to decrement to zero or to a DC value representing zerooutput. The combination of zero fader gain and zero integrator outputeffectively disables TRC mode attitude controller 408. When the groundspeed of the rotorcraft exceeds the 10 kts threshold, the gain of fader432 linearly decreases to zero over a predetermined time period or achange in predetermined speed range and the input of the integrator isswitched from the output of integral gain block 438 to the output oflimiter 442. Once the input to integrator is switched to the output oflimiter 442 via switch 452, the feedback action of the loop formed byintegrator 440, feedback gain 444 and limiter 442 forces the output ofintegrator 440 decrement to zero over a period of time. Limiter 442limits the rate at which integrator 440 is decremented such that thedecay of integrator 440 has more of a ramp response instead of anexponential response for high integrator output values. Thus, in someembodiments, during a mode transition from TRC mode to rate controlmode, the actuator command generated by summer 446 produces a gentlydecreasing or decaying signal from the forward path of the TRC modeattitude controller 408 due to fader 432 that is summed with a gentlydecreasing or decaying integrator value produced by integrator 440. Atthe same time as the output of the TRC mode attitude controller 408 isdecaying, the pilot stick command to rate controller 315 is increasingdue the increasing gain of fader 404. The net result of the decreasingoutput of the TRC mode attitude controller 408 and the increasing outputof fader 404 a or 404 b to rate controller 315 is a smooth transitionfrom TRC mode to a rate mode in some embodiments. In variousembodiments, the time that it takes for faders 404 a, 404 b and 432 tochange their gains may be the same or may be different from each other.In some embodiments, the fade-in and fade-out times of faders 404 a, 404b and 432 may be the same or they may be different from each otherdepending on the change in speed.

Similarly, during a mode transition from rate mode to TRC mode, thepilot stick control signal produced by fader 404 a or 404 b decreaseswhile the forward path of TRC mode attitude controller 408 increases dueto the gain of fader 432 increasing and due to the integral path of PIcontroller 450 being activated by coupling the input of integrator 440to the output of integral gain block 438 via switch 452.

In alternative embodiments of the present invention, integrator may bedecremented in a different manner. For example, instead of usingfeedback gain 444 to decrease the output of the integrator, a fader maybe coupled to the output of integrator 440 or may be coupled aftersummer 446.

In various embodiments, the transfer function and implementation ofdynamic compensation block 434, the gains of fader 432, proportionalgain 435, integral gain 438 and feedback gain 444, and the static anddynamic behavior of fader 432, limiter 442 and integrator 440 may beadjusted according to the specification and requirements of therotorcraft and may have different implementations and values with regardto the roll, pitch and other attitude channels.

It should be understood that the transition between TRC and rate controlmode with respect to the various rate commands is just one example ofmany possible system configurations that utilize embodiment modeswitching systems and methods. In alternative embodiments, other controlchannels besides the pitch rate and roll rate control channel may becontrolled. For example, embodiments of the present invention may beused to transition between attitude hold/command mode and a speedhold/command mode, transition between attitude hold/command mode and aposition hold/command mode, transition between direct verticalcollective control mode and an altitude hold/command mode, and/ortransition between vertical speed hold/command mode and an altitudehold/command mode. It should be appreciated that these are only a fewexamples of many possible applications.

FIG. 6 illustrates a block diagram of an embodiment method 500 ofoperating a rotorcraft that may be executed, for example, by a flightcomputer of a fly-by-wire system. In an embodiment, the pilot stickcommand of the cyclic control is interpreted as a translational ratecommand in step 518 in which the flight computer controls the rotorcraftto have a translational velocity that is proportional to a physicaloffset of the cyclic control. During step 518, a dynamic controlleradjusts the attitude of the rotorcraft in order to maintain the desiredtranslational velocity. When the velocity of the rotorcraft exceeds athreshold V_(threshold2) according to step 520 or when the pilot stickcommand exceeds V_(threshold2), the operational mode of the flightcomputer transitions to a rate command mode in step 508 in which thepilot stick command of the cyclic control is interpreted as a ratecommand. In this mode, the flight computer controls the rotorcraft tohave a rate that is proportional to the physical offset of the cycliccontrol.

In order to provide a smooth transition from the translational commandmode of step 518 to the rate command mode of step 508 and to reduce theoccurrence of physical transients of the rotorcraft, steps 502, 504 and506 are executed by the flight computer. In step 502, the gain of afirst dynamic controller, which may be implemented, for example, asouter loop controller that provides a rate control in response to adesired attitude command, is faded out over first period of time. Instep 504, a state of an integrator of the first dynamic controller isdecreases over a second period of time, and in step 506 a gain of thepilot stick command is faded in over a third period of time. In someembodiments, steps 502, 504 and 506 occur concurrently such that thegain of the first dynamic controller is reduced at the same time as thegain of the pilot stick command to the rate controller is faded in.During this time, the value of the integrator is reduced. In someembodiments, the value of the integrator is decremented by disconnectingthe integrator from the input of the first dynamic controller andcoupling the input of the integrator to its output. The input and/oroutput value of the integrator may be limited while the integrator valueis reduced in order to increase the amount of time that it takes todecrement the value of the integrator to zero.

The rate command continues to be determined by the pilot stick commandin step 508 until the velocity of the rotorcraft is less than or equalto threshold V_(threshold1) and the stick command S_(command) is lessthan an offset S_(threshold) that represents a particular translationalvelocity command according to step 510. In some embodiments,S_(threshold) may be set to a pilot stick offset that is less than 3kts. Alternatively, other thresholds may be used. Once this condition isdetected, the operational mode of the flight computer transitions to thetranslational rate command mode of step 518 via steps 512, 514 and 516.In some embodiments, threshold V_(threshold1) is equal to thresholdV_(threshold2), while in other embodiments, V_(threshold2) is greaterthreshold V_(threshold1) in order to provide hysteresis. In step 512,the gain of the first dynamic controller is faded in, in step 514, theintegrator of the first dynamic controller is reactivated by couplingits input to the signal path of the first dynamic controller, and instep 516 the gain of the pilot stick command to the rate controller isfaded out. In some embodiments, steps 512, 514 and 516 occurconcurrently such that the gain of the first dynamic controller isincreased at the same time as the gain of the pilot stick command to therate controller is reduced and at the same time the integrator of thefirst dynamic controller is reactivated.

FIG. 7 illustrates a computer system 601. The computer system 601 can beconfigured for performing one or more functions with regard to theoperation of the flight control system 201 and methods 500 and 1100, asdescribed herein. Further, any processing and analysis can be partly orfully performed by the computer system 601. The computer system 601 canbe partly or fully integrated with other aircraft computer systems orcan be partly or fully removed from the rotorcraft.

The computer system 601 can include an input/output (I/O) interface 603,an analysis engine 605, and a database 607. Alternative embodiments cancombine or distribute the I/O interface 603, the analysis engine 605,and the database 607, as desired. Embodiments of the computer system 601may include one or more computers that include one or more processorsand memories configured for performing tasks described herein. This caninclude, for example, a computer having a central processing unit (CPU)and non-volatile memory that stores software instructions forinstructing the CPU to perform at least some of the tasks describedherein. This can also include, for example, two or more computers thatare in communication via a computer network, where one or more of thecomputers include a CPU and non-volatile memory, and one or more of thecomputer's non-volatile memory stores software instructions forinstructing any of the CPU(s) to perform any of the tasks describedherein. Thus, while the exemplary embodiment is described in terms of adiscrete machine, it should be appreciated that this description isnon-limiting, and that the present description applies equally tonumerous other arrangements involving one or more machines performingtasks distributed in any way among the one or more machines. It shouldalso be appreciated that such machines need not be dedicated toperforming tasks described herein, but instead can be multi-purposemachines, for example computer workstations, that are suitable for alsoperforming other tasks.

The I/O interface 603 can provide a communication link between externalusers, systems, and data sources and components of the computer system601. The I/O interface 603 can be configured for allowing one or moreusers to input information to the computer system 601 via any knowninput device. Examples can include a keyboard, mouse, touch screen,and/or any other desired input device. The I/O interface 603 can beconfigured for allowing one or more users to receive information outputfrom the computer system 601 via any known output device. Examples caninclude a display monitor, a printer, cockpit display, and/or any otherdesired output device. The I/O interface 603 can be configured forallowing other systems to communicate with the computer system 60 i. Forexample, the I/O interface 603 can allow one or more remote computer(s)to access information, input information, and/or remotely instruct thecomputer system 601 to perform one or more of the tasks describedherein. The I/O interface 603 can be configured for allowingcommunication with one or more remote data sources. For example, the I/Ointerface 603 can allow one or more remote data source(s) to accessinformation, input information, and/or remotely instruct the computersystem 601 to perform one or more of the tasks described herein.

The database 607 provides persistent data storage for the computersystem 601. Although the term “database” is primarily used, a memory orother suitable data storage arrangement may provide the functionality ofthe database 607. In alternative embodiments, the database 607 can beintegral to or separate from the computer system 601 and can operate onone or more computers. The database 607 preferably provides non-volatiledata storage for any information suitable to support the operation ofthe flight control system 201 and the method 500, including varioustypes of data discussed further herein. The analysis engine 605 caninclude various combinations of one or more processors, memories, andsoftware components.

Embodiment mode transition systems and methods can be applied to a widevariety of different flight mode transitions. For example, embodimentsof the present invention can be applied to mode change from a flightmode that provides attitude and/or speed control to a flight mode thatprovides rate control and vice-versa. Examples of various flight modesare shown in the table of FIG. 8A, and various mode transitions betweenthese modes are shown in the table of FIG. 8B. It should be understoodthat the modes and mode transitions detailed in FIGS. 8A and 8B are anon-exhaustive list of possible mode transitions. In alternativeembodiment of the present invention, other modes and mode transition maybe implemented depending on the particular system and itsspecifications.

As shown, in FIG. 8A, various flight modes may include the translationalrate command mode (TRC) as discussed above, an attitude command returnto trim speed (ACRTT) mode, an attitude command speed hold (ACSH) mode,an attitude command attitude hold (ACAH) mode, a rate command return totrip speed (RCRTT) mode, a rate command speed hold (RCSH) mode and arate commend attitude hold (RCAH) mode.

FIG. 9 illustrates a flight control system 900 according to anembodiment of the present invention. Flight control system 900 issimilar to flight control system 400 described above with respect toFIG. 4 , with the exception that TRC mode attitude controllers 408 a and408 b are replaced by outer loop attitude controllers 908 a and 908 b.During modes in which outer loop 313 issues roll attitude commands andpitch attitude commands, outer loop attitude controller 908 a applies adynamic control algorithm to the pitch attitude error to produce anouter loop pitch rate command, and outer loop attitude controller 408 bapplies a dynamic control algorithm to the roll attitude error toproduce an outer loop roll rate command. These outer loop rate commandsare applied to decoupler 319 via summing blocks 412 a and 412 b. Invarious embodiments, outer loop attitude controllers 908 a and 908 b areactive in the TRC, ACRTT, ACSH and ACAH mentioned above.

In the TRC mode, the outer loop controller 313 receives a command fromthe pilot stick represented by pilot stick pitch block 401 a and pilotstick roll block 401 b, determines a corresponding translational speedbased on the pilot stick command, and determines corresponding pitch androll attitude commands based on the determined translational speed.Outer loop attitude controllers 908 a and 908 b produce outer loop pitchrate and roll rate commands based on the pitch and roll attitudecommands produced by outer loop 313.

In the ACRTT mode, the pilot stick controls the attitude of therotorcraft when the pilot stick is out of detent. When the pilot stickis released and goes back into detent, the flight control systemmaintains the rotorcraft at a predetermined trim speed. During the ACRTTmode, the outer loop controller 313 receives a command from the pilotstick represented by pilot stick pitch block 401 a and pilot stick rollblock 401 b when the pilot stick is out of detent, determines acorresponding attitude based on the pilot stick command, and determinescorresponding pitch and roll attitude commands based on the determinedcommanded attitude. Outer loop attitude controllers 908 a and 908 bproduce outer loop pitch rate and roll rate commands based on the pitchand roll attitude commands produced by outer loop 313 to effect thecommanded attitude.

When the pilot stick is in detent in the ACRTT mode, outer loopcontroller 313 determines pitch and roll attitude commands thatcorrespond to the trim speed. In some embodiments, this trim speed isthe speed of the rotorcraft prior to the pilot stick being taken out ofdetent. Outer loop attitude controllers 908 a and 908 b produce outerloop pitch rate and roll rate commands based on the pitch and rollattitude commands produced by outer loop 313 to effect the determinedtrim speed.

In the ACSH mode, the pilot stick controls the attitude of therotorcraft when the pilot stick is out of detent. When the pilot stickis released and goes back into detent, the flight control systemmaintains the speed at which the rotorcraft was travelling immediatelyprior to the pilot stick going back into detent. During the ACSH mode,the outer loop controller 313 receives a command from the pilot stickrepresented by pilot stick pitch block 401 a and pilot stick roll block401 b when the pilot stick is out of detent, determines a correspondingattitude based on the pilot stick command, and determines correspondingpitch and roll attitude commands based on the determined commandedattitude. Outer loop attitude controllers 908 a and 908 b produce outerloop pitch rate and roll rate commands based on the pitch and rollattitude commands produced by outer loop 313 to effect the commandedattitude.

When the pilot stick is released and falls back into detent in the ACSHmode, the outer loop continues to issue pitch and roll attitude commandsto maintain the speed at which the rotorcraft was travelling immediatelyprior to the pilot stick being released.

In the ACAH mode, the pilot stick controls the attitude of therotorcraft when the pilot stick is out of detent. When the pilot stickis released and goes back into detent, the flight control system bringsthe aircraft back to level attitude. During the ACAH mode, the outerloop controller 313 receives a command from the pilot stick representedby pilot stick pitch block 401 a and pilot stick roll block 401 b whenthe pilot stick is out of detent, determines a corresponding attitudebased on the pilot stick command, and determines corresponding pitch androll attitude commands based on the determined commanded attitude. Outerloop attitude controllers 908 a and 908 b produce outer loop pitch rateand roll rate commands based on the pitch and roll attitude commandsproduced by outer loop 313 to effect the commanded attitude.

When the pilot stick is released and falls back into detent in the ACRTTmode, the outer loop continues to issue pitch and roll attitude commandsto maintain the speed the rotorcraft was travelling immediately prior tothe pilot stick being released.

In the RCRTT mode, the pilot stick controls the rate of the rotorcraftwhen the pilot stick is out of detent. When the pilot stick is releasedand goes back into detent, the flight control system maintains therotorcraft at a predetermined trim speed. During the RCRTT mode, pitchrate controller 315 a and roll rate controller 315 b receives commandsrepresented by pilot stick pitch block 401 a and pilot stick roll block401 b when the pilot stick is out of detent, while directly affectingthe rate of the rotorcraft.

When the pilot stick is in detent in the RCRTT mode, outer loopcontroller 313 determines pitch and roll attitude commands thatcorrespond to the predetermined trim speed. In some embodiments, thepredetermined trim speed is the speed of the rotorcraft prior to thepilot stick being taken out of detent. Outer loop attitude controllers908 a and 908 b produce outer loop pitch rate and roll rate commandsbased on the pitch and roll attitude commands produced by outer loopcontroller 313.

In the RCSH mode, the pilot stick controls the rate of the rotorcraftwhen the pilot stick is out of detent. When the pilot stick is releasedand goes back into detent, the flight control system maintains the speedat which the rotorcraft was travelling immediately prior to the pilotstick going back into detent. During the RCSH mode, pitch ratecontroller 315 a and roll rate controller 315 b, receives commandsrepresented by pilot stick pitch block 401 a and pilot stick roll block401 b when the pilot stick is out of detent, while directly affectingthe rate of the rotorcraft.

When the pilot stick goes into detent in the RCSH mode, outer loopcontroller 313 determines pitch and roll attitude commands thatcorrespond to speed at which the rotorcraft was travelling immediatelyprior to the pilot stick being released. Outer loop attitude controllers908 a and 908 b produce outer loop pitch rate and roll rate commandsbased on the pitch and roll attitude commands produced by outer loopcontroller 313.

In the RCAH mode, the pilot stick controls the rate of the rotorcraftwhen the pilot stick is out of detent. When the pilot stick is releasedand goes back into detent, the flight control system maintains theattitude of the rotorcraft immediately prior to the pilot stick goingback into detent. During the RCAH mode, pitch rate controller 315 a androll rate controller 315 b receives commands represented by pilot stickpitch block 401 a and pilot stick roll block 401 b when the pilot stickis out of detent, while directly affect the rate of the rotorcraft.

When the pilot stick goes into detent in the RCAH mode, outer loopcontroller 313 determines pitch and roll attitude commands thatcorrespond to the attitude of the rotorcraft immediately prior to thepilot stick going back in detent. Outer loop attitude controllers 908 aand 908 b produce outer loop pitch rate and roll rate commands based onthe pitch and roll attitude commands produced by outer loop controller313.

Embodiment fading techniques may be applied to transitions between pilotstick effected rate control and outer loop effected attitude/speedcontrol. These transitions may occur both within the same control modeor may occur between control modes. For example, in some embodiments,when the pilot stick goes into detent for the RCRTT, RCSH, and RCAHmodes, outer loop pitch and roll rate commands generated by pilot sticks401 a and 401 b are faded out using faders 904 a and 904 b, while theouter loop pitch rate and roll rate commands are faded in using outerloop controllers 908 a and 908 b using embodiment systems and methodsdescribed herein. On the other hand, when the pilot stick goes out ofdetent, outer loop pitch and roll rate commands generated by pilotsticks 401 a and 401 b are faded in using faders 904 a and 904 b, whilethe outer loop pitch rate and roll rate commands are faded out usingouter loop controllers 908 a and 908 b as using embodiment systems andmethods described herein.

FIG. 8B illustrates a table showing various possible transitions thatmay occur between the various embodiment control modes described abovewith respect to FIG. 8A. In various embodiments, the flight controlsystem may be configured to transition between control modes dependingon the speed at which the rotorcraft is traveling. Alternatively, othercriteria may be used in conjunction with or instead of the speed of therotorcraft. For example, the pilot stick position, stick rate, or logictoggle switch may be used to determine the present control mode. Itshould also be understood that the 21 mode changes scenarios depicted inFIG. 8B represents a non-exhaustive set of examples. In alternativeembodiments, other mode change scenarios that are not depicted in FIG.8B may be implemented according to embodiments of the present invention.

The table of FIG. 8A categories each mode change scenario in terms of aflight control mode that is active at speeds of X kts and below, acorresponding flight control mode that is active at speeds of beyond Xkts. The speed at which the mode transitions occurs may be selectedaccording to the particular rotorcraft and its specifications. In oneexample, the transition speed is 10 kts.

An “X” in the column entitled “Attitude to Rate Transition” indicatesthat embodiment fading systems and methods are used to fade in and outpilot-stick based attitude commands and outer loop generated attitudecommands using embodiment systems and methods described herein. Forexample, since the TRC mode relies on outer loop controller 313 togenerate attitude commands, while the RCRTT mode relies on pilot stickinput to generate pitch rate and roll rate commands, transitioning fromthe TRC mode to the RCRTT mode, embodiment mode transition techniquesare used to fade out the pitch and roll rate commands generated by outerloop attitude controllers 908 a and 908 b and fade in the pilot stickcommands via faders 904 a and 904 b. When transitioning from the RCRTTmode back to the TRC mode, embodiment mode transition techniques areused to fade in the pitch and roll rate commands generated by outer loopattitude controllers 908 a and 908 b and fade out the pilot stickcommands via faders 904 a and 904 b.

As shown in the table of FIG. 8B, mode transitions between TRC andRCRTT; TRC and RCSH; TRC and RCAH; ACRTT and RCRTT; ACRTT and RCRTT;ACRTT and RCAH; ACSH and RCRTT; ACSH and RCSH; ACSH and RCAH; ACAH andRCRTT; ACAH and RCSH; ACAH and RCAH; RCRTT and RCSH; RCRTT and RCAH andRCSH and RCAH may use embodiment fading techniques. On the other hand,mode transitions between TRC and ACRTT; TRC and ACSH; TRC and ACAH;ACRTT and ACSH; ACRTT and ACAH; and ACSH and ACAH involve modes in whichouter loop controller 313 generates pitch and roll attitude commands.Thus, the mode transition may be implemented in a way that does notinvolve fading between signal paths that include faders 904 a and 904 band signal paths that include outer loop attitude controllers 908 a and908 b.

FIG. 10 illustrates a block diagram of outer loop attitude controller908 according to an embodiment of the present invention that may be usedto implement outer loop attitude controllers 908 a and 908 b illustratedin FIG. 9 . Outer loop attitude controller 908 is implemented andoperates in a similar fashion as TRC mode attitude controller 408described hereinabove with respect to FIG. 5 . However, outer loopattitude controller 908 may be active in ACRTT, ACSH, ACAH modes andwhen the speed of the rotorcraft is at or below X kts.

FIG. 11 illustrates a block diagram of an embodiment method 1100 ofoperating a rotorcraft that may be executed, for example, by a flightcomputer of a fly-by-wire system. In an embodiment, the pilot stickcommand of the cyclic control is interpreted as an attitude or speedcommand in step 1118 in which the flight computer controls therotorcraft to have a translational velocity or attitude that isproportional to a physical offset of the cyclic control. This may occur,for example, in modes TRC, ACRTT, ACSH and ACAH as described above.During step 1118, a dynamic controller adjusts the attitude or speed ofthe rotorcraft in order to maintain the desired translational velocityor attitude. When the velocity or attitude of the rotorcraft exceeds athreshold V_(threshold2) according to step 1120 or when the pilot stickcommand exceeds V_(threshold2), the operational mode of the flightcomputer transitions to a rate command mode in step 1108 in which thepilot stick command of the cyclic control is interpreted as a ratecommand. This may occur, for example, in a transition to the RCRTT, RCSHand RCAH modes. In these modes, the flight computer controls therotorcraft to have a rate that is proportional to the physical offset ofthe cyclic control.

In order to provide a smooth transition from the speed/attitude commandmodes of step 1118 to the rate command mode of step 1108 and to reducethe occurrence of physical transients of the rotorcraft, steps 1102,1104 and 1106 are executed by the flight computer. In step 1102, thegain of a first dynamic controller, which may be implemented, forexample, as outer loop controller that provides a rate control inresponse to a desired attitude command, is faded out over first periodof time or given range of speed. In step 1104, a state of an integratorof the first dynamic controller is decreases over a second period oftime, and in step 1106 a gain of the pilot stick command is faded inover a third period of time. In some embodiments, steps 1102, 1104 and1106 occur concurrently such that the gain of the first dynamiccontroller is reduced at the same time as the gain of the pilot stickcommand to the rate controller is faded in. During this time, the valueof the integrator is reduced. In some embodiments, the value of theintegrator is decremented by disconnecting the integrator from the inputof the first dynamic controller and coupling the input of the integratorto its output. The input and/or output value of the integrator may belimited while the integrator value is reduced in order to increase theamount of time that it takes to decrement the value of the integrator tozero.

The rate command continues to be determined by the pilot stick commandin step 1108 until the velocity of the rotorcraft is less than or equalto threshold V_(threshold1) and the stick command S_(command) is lessthan an offset S_(threshold) that represents a particular translationalvelocity or attitude command according to step 1110. In someembodiments, S_(threshold) may be set to a pilot stick offset that isless than 3 kts. Alternatively, other thresholds may be used. Once thiscondition is detected, the operational mode of the flight computertransitions to the translational rate command mode of step 1118 viasteps 1112, 1114 and 1116. In some embodiments, threshold V_(threshold1)is equal to threshold V_(threshold2), while in other embodiments,V_(threshold2) is greater threshold V_(threshold1) in order to providehysteresis. In step 1112, the gain of the first dynamic controller isfaded in, in step 1114, the integrator of the first dynamic controlleris reactivated by coupling its input to the signal path of the firstdynamic controller, and in step 1116 the gain of the pilot stick commandto the rate controller is faded out. In some embodiments, steps 1112,1114 and 1116 occur concurrently such that the gain of the first dynamiccontroller is increased at the same time as the gain of the pilot stickcommand to the rate controller is reduced and at the same time theintegrator of the first dynamic controller is reactivated.

FIG. 12 illustrates a flight control system 1200 according to anembodiment of the present invention. Flight control system 1200 issimilar to flight control system 900 described above with respect toFIG. 9 , with the exception that the output of outer loop attitudecontrollers 908 a and 908 b are replaced by outer loop attitudecontrollers 1208 a and 1208 b, pitch rate controller 315 a is replace bypitch rate controller 1215 a, and roll rate controller 315 b is replacedby roll rate controller 1215 b, and summing blocks 412 a and 412 b areremoved. Instead of routing the outputs of outer loop attitudecontrollers 1208 a and 1208 b to summing blocks coupled to the outputsof pitch rate controller 1215 a and 1215 b, the outputs of outer loopattitude controllers 1208 a and 1208 b are coupled to inputs of pitchrate controller 1215 a and roll rate controller 1215 b, respectively.The functionality of flight control system 1200 shown in FIG. 12 may bethe same as the functionality of flight control system 900 shown in FIG.9 in some embodiments.

FIG. 13 illustrates a block diagram of outer loop attitude controller1208 that may be used to implement outer loop attitude controllers 1208a and 1208 b illustrated in FIG. 12 . As shown, outer loop attitudecontroller 1208 includes subtractor 430, fader 432, dynamic compensationblock 434 and proportional gain block 436. Subtractor 430 produces anerror signal based on the attitude command produced by the outer loopcontroller 313 (FIG. 12 ) and dynamic compensation block 434 compensatesfor the dynamics of the rotorcraft in order to improve stability and/oradjust the time response of the attitude loop. Dynamic compensationblock 434 may include various control blocks known in the art includingbut not limited a PID controller and a lead-lag compensator.Proportional gain block 436 is used to produce the rate command and maybe used to impart a proportional gain to the output of dynamiccompensation block 434. Proportional gain block 436 may be implemented,for example, by performing a multiplication or scaling operation.

During a first mode of operation when the rotorcraft has a ground speedbelow a predetermined speed threshold (represented by X kts), fader 432has a gain of one or some other constant gain. This effectively couplesthe output of outer loop controller 313 to one of rate controllers 1215a or 1215 b. On the other hand, in a second mode of operation in whichthe rotorcraft has a ground speed of X kts and above or the speedcommand by displacement on the cyclic controller is greater than X kts,fader 432 has a gain of zero. When fader 432 has a gain of zero, theoutput of outer loop controller 313 is effectively disconnected fromrate controller 1215 a and/or rate controller 1215 b.

In various embodiments, when the ground speed of the rotorcraft exceedsthe X kts threshold, the gain of fader 432 linearly decreases to zeroover a predetermined time period. In some embodiments, the speedthreshold is about 10 kts; however, other speed thresholds may be useddepending on the particular embodiment and its specifications. The firstand second modes of operation may represent any of the modes detailedabove with respect to FIGS. 8A and 8B. In some embodiments the firstmode of operation represents one of the modes of operation shown in thesecond column of FIG. 8B when the ground speed of the rotorcraft is Xkts and below. This first mode of operation may include, for example,TRC, ACRTT, ACSH, ACAH, RCRTT and/or RCSH. The second mode of operationrepresents one of the modes of operation shown in the third column ofFIG. 8B when the ground speed of the rotorcraft is greater than X kts.This second mode of operation may include, for example, ACAH, ACRTT,ACSH, RCAH, RCRTT and RCSH. In some embodiments, the first mode ofoperation is a speed control mode or an attitude control mode, while thesecond mode of operation is a rate control mode.

In some embodiments, mode switching systems and methods may be appliedto other types of aircraft besides rotorcraft. For example, in anembodiment flight control system directed to fixed-wing aircraft, achange in the pilot stick command may be configured to produce acorresponding change in speed or attitude in the first mode, while achange in the pilot stick command in the second mode of operation may beconfigured to produce a change in rate or to directly control thecontrol surfaces or the motor power in a fixed wing aircraft.

FIG. 14A illustrates a block diagram of rate controller 1215 that may beused to implement rate controllers 1215 a and 1215 b shown in FIG. 12 .As shown, rate controller 1215 includes summing block 1240, followed bydynamic compensation block 1244 and PID controller 1252. Summing block1240 forms a sum of the pilot stick rate command generated by pilotstick 401 a and/or 401 b and the output of outer loop attitudecontroller 1208 a and/or 1208 b. Dynamic compensation block 1244 maycompensate for the dynamics of the rotorcraft in order to improvestability and/or adjust the time response of the rate loop. Dynamiccompensation block 1244 may include various control blocks known in theart including but not limited a PID controller and a lead-lagcompensator. PID controller 1252 is used to produce the actuator commandand may be implemented using proportional gain block 1222, integrator1224 and differentiator 1228. In some embodiments PID controller 1252may be replaced by other dynamic controller topologies known in the art.Proportional gain block 1222 may be implemented, for example, byperforming a multiplication or scaling operation. Integrator 1224 may beimplemented, for example, using an accumulator; and differentiator 1228may be implemented, for example, using subtraction circuit that isconfigured to subtract a previous value from a present value. In someembodiments, the embodiments of FIGS. 12, 13 and 14A may be operatedaccording to method 1100 described above with respect to FIG. 11 .

FIG. 14B illustrates a block diagram of rate controller 1270 that can beused to implement rate controllers 1215 a and 1215 b according to analternative embodiment of the present invention. The structure of ratecontroller 1270 is similar to the structure of rate controller 1215shown in FIG. 14A, with the exception that the integrator of PIDcontroller 1220 can be activated and deactivated during operation in asimilar manner as integrator 440 in FIG. 10 described above.

During operation in the first mode when the rotorcraft has a groundspeed of less than X kts, the input of integrator 1224 is coupled to theoutput of dynamic compensation block 1244 via integral gain block 1234.On the other hand, in the second mode in which the rotorcraft has aground speed of X kts and above and the speed command by displacement onthe cyclic controller is greater than X kts, the input of integrator1224 is coupled to its output via feedback gain block 1226 and limiter1230, which effectively causes integrator 1224 to decrement to zero orto a DC value representing zero output. When the ground speed of therotorcraft exceeds the X kts threshold, the input of the integrator 1224is switched from the output of integral gain block 1234 to the output oflimiter 1230. Once the input to integrator 1224 is switched to theoutput of limiter 1230 via switch 1232, the feedback action of the loopformed by integrator 1224, feedback gain 1226 and limiter 1230 forcesthe output of integrator 1224 decrement to zero over a period of time.Limiter 1230 limits the rate at which integrator 1224 is decrementedsuch that the decay of integrator 1224 has more of a ramp responseinstead of an exponential response for high integrator output values. Insome embodiments, the fade-in and fade-out times of faders 904 a, 904 b,432, and 1230 may be the same or they may be different from each otherdepending on the change in speed.

FIG. 15 illustrates a block diagram of an embodiment method 1500 ofoperating an aircraft that may be executed, for example, by a flightcomputer of a fly-by-wire system. In some embodiments, method 1500 maybe applied to the embodiments described above with respect to FIGS. 12,13 and 14B. In an embodiment, the pilot stick command is interpreted asan attitude or speed command in step 1518 in which the flight computercontrols the aircraft to have a translational velocity or attitude thatis proportional to a physical offset of the pilot stick. This may occur,for example, in modes TRC, ACRTT, ACSH and ACAH as described above.During step 1518, an attitude controller adjusts the attitude or speedof the aircraft in order to maintain the desired translational velocityor attitude. When the velocity or attitude of the rotorcraft exceeds athreshold V_(threshold2) according to step 1520 or when the pilot stickcommand exceeds V_(threshold2), the operational mode of the flightcomputer transitions to a rate command mode in step 1508 in which thepilot stick command is interpreted as a rate command. This may occur,for example, in a transition to the RCRTT, RCSH and RCAH modes. In thesemodes, the flight computer controls the rotorcraft to have a rate thatis proportional to the physical offset of the pilot stick.

In order to provide a smooth transition from the speed/attitude commandmodes of step 1518 to the rate command mode of step 1508 and to reducethe occurrence of physical transients of the rotorcraft, steps 1502 and1506 are executed by the flight computer. In step 1502, the gain of anattitude controller, which may be implemented, for example, as outerloop controller that provides a response to a desired attitude command,is faded out over first period of time or given range of speed. In step1506 a gain of the pilot stick command is faded in over a second periodof time. In some embodiments, steps 1502 and 1506 occur concurrentlysuch that the gain of the attitude controller is reduced at the sametime as the gain of the pilot stick command to the rate controller isfaded in.

The rate command continues to be determined by the pilot stick commandin step 1508 until the velocity of the rotorcraft is less than or equalto threshold V_(threshold1) and the pilot stick command S_(command) isless than an offset S_(threshold) that represents a particulartranslational velocity or attitude command according to step 1510. Insome embodiments, S_(threshold) may be set to a pilot stick offset thatis less than 3 kts. Alternatively, other thresholds may be used. Oncethis condition is detected, the operational mode of the flight computertransitions to the translational rate command mode of step 1518 viasteps 1512 and 1516. In some embodiments, threshold V_(threshold1) isequal to threshold V_(threshold2), while in other embodiments,V_(threshold2) is greater threshold V_(threshold1) in order to providehysteresis. In step 1512, the gain of the attitude controller is fadedin, and in step 1516 the gain of the pilot stick command to the ratecontroller is faded out. In some embodiments, steps 1512 and 1516 occurconcurrently such that the gain of the attitude controller is increasedat the same time as the gain of the pilot stick command to the ratecontroller is reduced.

Embodiments of the present invention are summarized here. Otherembodiments can also be understood from the entirety of thespecification and the claims filed herein.

Example 1. A method of operating an aircraft includes operating theaircraft in a first mode including determining an attitude based on apilot stick signal generated by a pilot stick assembly, determining afirst rate command based on the determined attitude using an attitudecontroller, determining an actuator command based on the determinedfirst rate command, determining the actuator command including using arate controller having an integrator, and providing an output of therate controller to an actuator, where a translational speed or anattitude of the aircraft is proportional to an amplitude of the pilotstick signal in the first mode; transitioning from the first mode to asecond mode when a velocity of the aircraft exceeds a first velocitythreshold, transitioning including fading out a gain of the attitudecontroller over a first period of time; and operating the aircraft inthe second mode including providing the pilot stick signal to an inputof the rate controller, where the output of the rate controller isproportional to the amplitude of the pilot stick signal.

Example 2. The method of example 1, where transitioning from the firstmode to the second mode further includes decreasing a value of theintegrator over a second period of time.

Example 3. The method of one of examples 1 or 2, where determining theattitude based on the pilot stick signal includes determining thetranslational speed based on the pilot stick signal by determining aforward speed component based on a cyclic longitude stick controlportion of the pilot stick signal and determining a lateral speedcomponent based on a cyclic latitude stick control portion of the pilotstick signal; and determining a pitch attitude based on the determinedforward speed component and a roll attitude based on the determinedlateral speed component.

Example 4. The method of one of examples 1 to 3, where determining theattitude based on the pilot stick signal includes determining a pitchattitude based on a cyclic longitude stick control portion of the pilotstick signal and determining a roll attitude based on a cyclic latitudestick control portion of the pilot stick signal.

Example 5. The method of one of examples 1 to 4, further includingtransitioning from the second mode to the first mode when the velocityof the aircraft decreases below a second velocity threshold,transitioning from the second mode to the first mode including: fadingin the gain of the attitude controller over a third period of time, andreactivating the integrator.

Example 6. The method of example 5, further including: fading in a gainof the pilot stick signal in the rate controller over a fourth period oftime when a pilot stick of the pilot stick assembly is out of detent;and fading out the gain of the pilot stick signal in the rate controllerover a fifth period of time when the pilot stick of the pilot stickassembly is in detent.

Example 7. The method of one of examples 5 or 6, where the firstvelocity threshold is greater than the second velocity threshold.

Example 8. The method of one of examples 1 to 7, where decreasing thevalue of the integrator includes coupling the output of the integratorto an input of the integrator via a first feedback path.

Example 9. The method of example 8, where the first feedback pathincludes a limiter.

Example 10. The method of one of examples 1 to 9, where the first modeincludes a translational rate command, where a speed of the aircraft isproportional to the pilot stick signal.

Example 11. The method of one of examples 1 to 10, where the first modeincludes an attitude command, where an attitude of the aircraft isproportional to the pilot stick signal.

Example 12. A flight control system for an aircraft including: aprocessor and a non-transitory computer readable storage medium with anexecutable program stored thereon, the executable program includinginstructions to: receive a pilot control signal via a first interface ofthe processor; in a first mode determine an attitude based on thereceived pilot control signal, determine a first rate command based onthe determined attitude using an attitude controller, determine anactuator command based on the determined first rate command, determiningthe actuator command including executing a rate controller that has anintegrator, and providing an output of the rate controller to anactuator via a second interface of the processor, where a state of theaircraft corresponding to the attitude is configured to be proportionalto the received pilot control signal; and transitioning from the firstmode to a second mode when a first condition of the aircraft crosses afirst predetermined threshold, transitioning including fading out a gainof the attitude controller over a first period of time, and decreasing avalue of the integrator over a second period of time; and in the secondmode, providing the pilot control signal to an input of the ratecontroller, where the output of the rate controller is proportional tothe received pilot control signal.

Example 13. The flight control system of example 12, where in the firstmode, the received pilot control signal represents a translational speedcommand, the determined attitude is based on the translational speedcommand, and the first condition of the aircraft is a velocity of theaircraft.

Example 14. The flight control system of one of examples 12 or 13, wherein the first mode, the received pilot control signal represents anattitude command, the determined attitude is based on the attitudecommand, and the first condition of the aircraft is a velocity of theaircraft.

Example 15. The flight control system of one of examples 12 to 14,where: the determined attitude includes a pitch attitude and a rollattitude; executing the rate controller includes executing a first pitchdynamic controller having a pitch integrator and executing a first rolldynamic controller having a roll integrator; and providing the output ofthe rate controller to the actuator via the second interface of theprocessor includes providing an output of the first pitch dynamiccontroller to a pitch actuator and providing an output of the first rolldynamic controller to a roll actuator.

Example 16. The flight control system of one of examples 12 to 15, wherethe executable program further includes instructions to fade in a gainof the pilot control signal in the rate controller when a pilot controlstick is out of detent; and fade out the gain of the pilot controlsignal in the rate controller when the pilot control stick is in detent.

Example 17. The flight control system of one of examples 12 to 16, wherethe executable program further includes instructions to transition fromthe second mode to the first mode when the first condition of theaircraft decreases below a second predetermined threshold, wheretransitioning from the second mode to the first mode includes: fading inthe gain of the attitude controller over a third period of time, andreactivating the integrator.

Example 18. The flight control system of example 17, where the firstpredetermined threshold is greater than the second predeterminedthreshold.

Example 19. The flight control system of one of examples 12 to 18, wheredecreasing the value of the integrator includes coupling the output ofthe integrator to an input of the integrator via a first feedback path.

Example 20. The flight control system of example 19, where the firstfeedback path includes a limiter.

Example 21. The flight control system of one of examples 12 to 20, wherethe first mode includes a translational rate command, where a speed ofthe aircraft is proportional to the pilot control signal.

Example 22. The flight control system of one of examples 12 to 21, wherethe first mode includes an attitude command, where an attitude of theaircraft is proportional to the pilot control signal.

Example 23. A rotorcraft including: a body; a power train coupled to thebody and including a power source and a drive shaft coupled to the powersource; a rotor system coupled to the power train and including aplurality of rotor blades; a flight control system operable to change atleast one operating condition of the rotor system; a pilot controlassembly configured to receive commands from a pilot, where the flightcontrol system is a fly-by-wire flight control system in electricalcommunication with the pilot control assembly; and a flight controlcomputer in electrical communication between the flight control systemand the pilot control assembly, the flight control computer configuredto: receive, from the pilot control assembly a pilot command to change afirst flight characteristic, when a velocity of the rotorcraft is lessthan a first velocity threshold, interpret the first flightcharacteristic as a requested translational speed or as a requestedattitude in a first mode, determine a controlled attitude based on therequested translational speed or the requested attitude using anattitude controller, determining a first rate command based on thedetermined attitude, and determine an actuator command based on thedetermined rate command using a rate controller; when the velocity ofthe rotorcraft is greater than a second velocity threshold, interpretthe first flight characteristic as a rate in a second mode by providingthe pilot command to the rate controller, and when the velocity of therotorcraft increases past the second velocity threshold, fade out a gainof the attitude controller, and successively decrease a value of anintegrator in the rate controller.

Example 24. The rotorcraft of example 23, where: the controlled attitudeincludes a controlled pitch attitude; and the rate includes a pitchrate.

Example 25. The rotorcraft of example 23, where: the controlled attitudefurther includes a controlled roll attitude; and the rate furtherincludes a roll rate.

Example 26. The rotorcraft of one of examples 23 to 25, where the flightcontrol computer is further configured to transition from the secondmode to the first mode when the velocity of the rotorcraft decreasesbelow the first velocity threshold, and transitioning from the secondmode to the first mode includes: fading in the gain of the attitudecontroller, and reactivating the integrator.

Example 27. The rotorcraft of one of examples 23 to 26, where the flightcontrol computer is further configured to: fade in a gain of the pilotcommand to the rate controller when a pilot stick of the pilot controlassembly is out of detent; and fade out the gain of the pilot command tothe rate controller when the pilot stick of the pilot control assemblyis in detent.

Advantages of embodiments include the ability to automatically andsmoothly transition between a TRC mode and a rate control mode orbetween a translational speed hold mode and a rotational rate controlmode. Further advantages include making the transition smooth andseamless enough that little change in attitude is noticeable to pilot oroccupants of the aircraft.

While this invention has been described with reference to illustrativeembodiments, this description is not intended to be construed in alimiting sense. Various modifications and combinations of theillustrative embodiments, as well as other embodiments of the invention,will be apparent to persons skilled in the art upon reference to thedescription. It is therefore intended that the appended claims encompassany such modifications or embodiments.

What is claimed is:
 1. A method of operating an aircraft, the methodcomprising: operating the aircraft in a first mode, the operating in thefirst mode comprising providing an attitude input signal to a ratecontroller having an integrator, wherein the attitude input signal isbased on a pilot control signal provided via an interface, providing anoutput of the rate controller to an actuator, wherein a translationalspeed or an attitude of the aircraft is proportional to an amplitude ofthe pilot control signal in the first mode; transitioning from the firstmode to a second mode when a velocity of the aircraft exceeds a firstvelocity threshold, the transitioning comprising: fading out a gainapplied to the attitude input signal over a first predetermined periodof time, and successively decreasing a value of the integrator over asecond predetermined period of time; and operating the aircraft in thesecond mode, the operating in the second mode comprising providing thepilot control signal to an input of the rate controller.
 2. The methodof claim 1, further comprising determining the attitude input signal,wherein determining the attitude input signal comprises: determining thetranslational speed based on the pilot control signal by determining aforward speed component based on a cyclic longitude control portion ofthe pilot control signal and determining a lateral speed component basedon a cyclic latitude control portion of the pilot control signal; anddetermining a pitch attitude based on the determined forward speedcomponent and a roll attitude based on the determined lateral speedcomponent.
 3. The method of claim 1, further comprising determining theattitude input signal, wherein determining the attitude input signalcomprises determining a pitch attitude based on a cyclic longitudecontrol portion of the pilot control signal and determining a rollattitude based on a cyclic latitude control portion of the pilot controlsignal.
 4. The method of claim 1, wherein the first mode comprises atranslational rate command mode, wherein a speed of the aircraft isproportional to the pilot control signal.
 5. The method of claim 1,wherein the first mode comprises an attitude command mode, wherein theattitude of the aircraft is proportional to the pilot control signal. 6.The method of claim 1, further comprising transitioning from the secondmode to the first mode when the velocity of the aircraft decreases belowa second velocity threshold, transitioning from the second mode to thefirst mode comprising: fading in the gain applied to the attitude inputsignal over a third predetermined period of time; and reactivating theintegrator.
 7. The method of claim 6, wherein the first velocitythreshold is greater than the second velocity threshold.
 8. The methodof claim 1, wherein decreasing the value of the integrator comprisescoupling an output of the integrator to an input of the integrator via afirst feedback path.
 9. The method of claim 8, wherein the firstfeedback path comprises a limiter.
 10. A flight control system for anaircraft comprising: a processor and a non-transitory computer readablestorage medium with an executable program stored thereon, the executableprogram including instructions to: receive a pilot control signal via afirst interface of the processor; in a first mode providing an attitudeinput signal to a rate controller having an integrator, wherein theattitude input signal is based on the pilot control signal; provide anoutput of the rate controller to an actuator via a second interface ofthe processor, wherein a state of the aircraft corresponding to anattitude of the aircraft is configured to be proportional to thereceived pilot control signal; and transition from the first mode to asecond mode when a first condition of the aircraft crosses a firstpredetermined threshold, the transition comprising: fading out a gainapplied to the attitude input signal over a first predetermined periodof time; decreasing a value of the integrator over a secondpredetermined period of time; and in the second mode, providing thepilot control signal to an input of the rate controller.
 11. The flightcontrol system of claim 10, wherein in the first mode, the receivedpilot control signal represents a translational speed command, theattitude input signal is based on the translational speed command, andthe first condition of the aircraft is a velocity of the aircraft. 12.The flight control system of claim 10, wherein in the first mode, thereceived pilot control signal represents an attitude command, theattitude input signal is based on the attitude command, and the firstcondition of the aircraft is a velocity of the aircraft.
 13. The flightcontrol system of claim 10, wherein the executable program furthercomprises instructions to execute a first pitch dynamic controllerhaving a pitch integrator and instructions to execute a first rolldynamic controller having a roll integrator, wherein providing theoutput of the rate controller to the actuator via the second interfaceof the processor comprises providing an output of the first pitchdynamic controller to a pitch actuator and providing an output of thefirst roll dynamic controller to a roll actuator.
 14. The flight controlsystem of claim 10, wherein the first mode comprises a translationalrate command mode, wherein a speed of the aircraft is proportional tothe pilot control signal.
 15. The flight control system of claim 10,wherein the first mode comprises an attitude command mode, wherein anattitude of the aircraft is proportional to the pilot control signal.16. The flight control system of claim 10, wherein the executableprogram further includes instructions to transition from the second modeto the first mode when the first condition of the aircraft decreasesbelow a second predetermined threshold, wherein transitioning from thesecond mode to the first mode comprises: fading in the gain applied tothe attitude input signal over a third predetermined period of time; andreactivating the integrator.
 17. The flight control system of claim 16,wherein the first predetermined threshold is greater than the secondpredetermined threshold.
 18. The flight control system of claim 10,wherein decreasing the value of the integrator comprises coupling theoutput of the integrator to an input of the integrator via a firstfeedback path.
 19. The flight control system of claim 18, wherein thefirst feedback path comprises a limiter.
 20. A method of operating anaircraft, the method comprising: operating the aircraft in a first mode,the operating in the first mode comprising: determining a translationalspeed based on a pilot control signal provided via an interface;determining an attitude based on the determined translational speed;determining an actuator command based on the determined attitude,determining the actuator command comprising using a first dynamiccontroller having an integrator; providing an output of the firstdynamic controller to an actuator, wherein a translational speed of theaircraft is proportional to an amplitude of the pilot control signal inthe first mode, and a change in the amplitude of the pilot controlsignal causes a change in the translational speed of the aircraft; andtransitioning from the first mode to a second mode when a velocity ofthe aircraft exceeds a first velocity threshold, the transitioningcomprising: fading out a gain of the first dynamic controller over afirst period of time; decreasing a value of the integrator over a secondperiod of time; and operating the aircraft in the second mode, theoperating in the second mode comprising providing an output of anattitude rate controller to the actuator.
 21. A method of operating anaircraft, the method comprising: operating the aircraft in a first mode,the operating in the first mode comprising: determining an attitudebased on a pilot control signal provided via an interface; determiningan actuator command based on the determined attitude, determining theactuator command comprising using a first dynamic controller having anintegrator; providing an output of the first dynamic controller to anactuator, wherein a translational speed or the attitude of the aircraftis proportional to an amplitude of the pilot control signal in the firstmode; transitioning from the first mode to a second mode when a velocityof the aircraft exceeds a first velocity threshold, the transitioningcomprising: fading out a gain of the first dynamic controller over afirst period of time; decreasing a value of the integrator over a secondperiod of time; and operating the aircraft in the second mode, theoperating in the second mode comprising providing an output of a ratecontroller to the actuator.